Hybrid Continuous Fiber Chopped Fiber Polymer Composite Structure

ABSTRACT

An engine component having a monolithic composite body, the body having a continuous fiber portion, a chopped fiber portion, a thermoplastic polymer contained in both the continuous fiber portion and the chopped fiber portion and between the continuous and chopped fiber portions.

CROSS-REFERENCE TO RELATED APPLICATIONS

None

BACKGROUND

The disclosed embodiments generally pertain to aircraft engine parts.

More particularly, but not by way of limitation, present embodiments relate to aircraft engine parts formed of hybrid composite materials to form more complicated geometries.

A typical gas turbine engine generally possesses a forward end and an aft end with its several core or propulsion components positioned axially therebetween. An air inlet or intake is at a forward end of the engine. Moving toward the aft end, in order, the intake is followed by a compressor, a combustion chamber, a turbine, and a nozzle at the aft end of the engine. It will be readily apparent from those skilled in the art that additional components may also be included in the engine, such as, for example, low-pressure and high-pressure compressors, and high-pressure and low-pressure turbines. This, however, is not an exhaustive list. An engine also typically has an internal shaft axially disposed along a center longitudinal axis of the engine. The internal shaft is connected to both the turbine and the air compressor, such that the turbine provides a rotational input to the air compressor to drive the compressor blades.

In operation, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases which flow downstream through turbine stages. These turbine stages extract energy from the combustion gases. A high pressure turbine first receives the hot combustion gases from the combustor and includes a stator nozzle assembly directing the combustion gases downstream through a row of high pressure turbine rotor blades extending radially outwardly from a supporting rotor disk. In a two stage turbine, a second stage stator nozzle assembly is positioned downstream of the first stage blades followed in turn by a row of second stage rotor blades extending radially outwardly from a second supporting rotor disk. The turbine converts the combustion gas energy to mechanical energy.

It is always desirable to reduce the weight of a gas turbine engine utilized in the aviation industry. Such weight reduction results in higher efficiency of the engine, which reduces costs for operators. In attempting to reach this goal, designers have turned to alternative materials in producing parts. In areas where temperatures are reduced, engine designers have attempted to utilize parts formed of polymer composite materials. This has led to at least two design issues to overcome.

First, designers desire to create parts or components that can withstand the rigors of high speed aircraft engine operation. Second, designers are limited by the shape and geometry of the parts being created, as related to strength and performance requirements for those parts. Complicated geometries are difficult to create from composites that have continuous fiber reinforcement, which is required for higher strength applications. In sum, it is difficult to manufacture aircraft engine composite components having high loading capabilities and complex shapes.

As may be seen by the foregoing, it would be desirable to overcome these and other deficiencies with gas turbine engines components.

SUMMARY

According to present aspects, a hybrid composite architecture is disclosed which enables the design and manufacturing of high-performance monolithic structural parts that have complex secondary features. Notable, but non-limiting, examples of such include spinner cones and aft support rings, although various other parts may be formed.

According to still other aspects of this disclosure, an engine component may be produced which exhibits mechanical, chemical and thermal properties (including strength, fatigue resistance, maximum temperature capability and chemical/fluid resistance) suitable for use in aircraft applications.

According to still other aspects of the disclosure, polymer composite aircraft parts may be formed being constructed of first portions formed of continuous fiber reinforcement and second portions formed of chopped fibers.

According to at least some embodiments, an aircraft engine component, comprises a monolithic composite body, the body having a continuous fiber portion, a chopped fiber portion, a thermoplastic polymer contained in both the continuous fiber portion and the chopped fiber portion and between the continuous and chopped fiber portions. The aircraft engine component wherein the thermoplastic polymer is one of PEEK, PPS, PEKK, and PEI. The aircraft engine component wherein the fiber is one of carbon fiber, glass fiber and a mixture of the carbon fiber and said glass fiber. The aircraft engine component wherein the continuous fiber is one a braid, woven fiber and a unidirectional tape. The aircraft engine component wherein the chopped fiber constituent is formed of a unidirectional preimpregnated tape. The aircraft engine component wherein the chopped fiber length is less than one inch (1″). The aircraft engine component further comprising one or more co-molded metallic features. The aircraft engine component wherein the co-molded metallic features are one of a flange, bushing and threaded insert. The aircraft engine component wherein the continuous fiber of the composite body carries a hoop load. The aircraft engine component wherein the continuous fiber is a braided preform that contains at least one of fibers of carbon, glass and thermoplastic. The aircraft engine component wherein the at least one of fibers is dry carbon fiber and thermoplastic fiber. The aircraft engine component wherein the at least one of fibers is dry carbon fiber, glass fiber and thermoplastic fiber. The aircraft engine component wherein the chopped fiber portion fails to carry a structural load. The aircraft engine component wherein the chopped fiber portion is adjacent to an engine air flowpath. The aircraft engine component further comprising an erosion protection layer on an outermost surface of said chopped fiber portion. The aircraft engine component wherein the component is a rotating part. The aircraft engine component wherein the rotating part is a spinner nose cone. The aircraft engine component wherein the rotating part being a spinner support ring.

All of the above outlined features are to be understood as exemplary only and many more features and objectives of the invention may be gleaned from the disclosure herein. Therefore, no limiting interpretation of this summary is to be understood without further reading of the entire specification, claims, and drawings included herewith.

BRIEF DESCRIPTION OF THE ILLUSTRATIONS

Embodiments of the invention are illustrated in the following illustrations.

FIG. 1 is a side section view of a gas turbine engine.

FIG. 2 is an isometric view of an exemplary composite component.

FIG. 3 is an isometric view of a second exemplary composite component.

FIG. 4 is a side section of the exemplary embodimentn including the component of FIG. 3.

FIG. 5 is a sectional view of the aerodynamic feature of the component of FIG. 4.

DETAILED DESCRIPTION

Reference now will be made in detail to embodiments provided, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation, not limitation of the disclosed embodiments. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present embodiments without departing from the scope or spirit of the disclosure. For instance, features illustrated or described as part of one embodiment can be used with another embodiment to still yield further embodiments. Thus it is intended that the present invention covers such modifications and variations as come within the scope of the appended claims and their equivalents.

Referring to FIGS. 1-5 various embodiments of structures constructed of polymer matrix composite (PMC) materials and processes are taught. More specifically, hybrid continuous fiber-chopped fiber polymer composite structures for aircraft engine applications are shown and described capable of use in a wide range of applications, for example, aircraft engine components, and more particularly, fan areas and by-pass portions of gas turbine engines. The hybrid polymer composite structures are suitable for use in a variety of locations and according to the non-limiting examples are utilized in areas wherein temperatures and loading requirements may be met through the use of the composite structures. The hybrid polymer composite structures are monolithic and may be formed of both continuous and chopped fibers wherein the continuous fibers may be laid for shapes which are of more simple geometric shape while more complicated geometries, not readily formable with continuous fiber composites, are formed with the chopped fiber composites. The term monolithic is utilized to mean that the same polymer is used in the continuous fiber reinforced section(s) and the chopped fiber reinforced section(s). As a result, the two fiber types are joined by polymers common to both fiber types, for example thermoplastic resin.

As used herein, the terms “axial” or “axially” refer to a dimension along a longitudinal axis of an engine. The term “forward” used in conjunction with “axial” or “axially” refers to moving in a direction toward the engine inlet, or a component being relatively closer to the engine inlet as compared to another component. The term “aft” used in conjunction with “axial” or “axially” refers to moving in a direction toward the engine nozzle, or a component being relatively closer to the engine nozzle as compared to another component.

As used herein, the terms “radial” or “radially” refer to a dimension extending between a center longitudinal axis of the engine and an outer engine circumference. The use of the terms “proximal” or “proximally,” either by themselves or in conjunction with the terms “radial” or “radially,” refers to moving in a direction toward the center longitudinal axis, or a component being relatively closer to the center longitudinal axis as compared to another component. The use of the terms “distal” or “distally,” either by themselves or in conjunction with the terms “radial” or “radially,” refers to moving in a direction toward the outer engine circumference, or a component being relatively closer to the outer engine circumference as compared to another component.

As used herein, the terms “lateral” or “laterally” refer to a dimension that is perpendicular to both the axial and radial dimensions.

Referring initially to FIG. 1, a schematic side section view of a gas turbine engine 10 is shown having an engine inlet end 12 wherein air enters the propulsor 13 which is defined generally by a compressor 14, a combustor 16 and a multi-stage high pressure turbine 20. Collectively, the propulsor 13 provides thrust or power during operation. The gas turbine 10 may be used for aviation, power generation, industrial, marine or the like.

In operation air enters through the air inlet end 12 of the engine 10 and moves through at least one stage of compression where the air pressure is increased and directed to the combustor 16. The compressed air is mixed with fuel and burned providing the hot combustion gas which exits the combustor 16 toward the high pressure turbine 20. At the high pressure turbine 20, energy is extracted from the hot combustion gas causing rotation of turbine blades which in turn cause rotation of the shaft 24. The shaft 24 passes toward the front of the engine to continue rotation of the one or more compressor stages 14, a turbofan 18 or inlet fan blades, depending on the turbine design. The turbofan 18 is connected by the shaft 28 to a low pressure turbine 21 and creates thrust for the turbine engine 10. A low pressure turbine 21 may also be utilized to extract further energy and power additional compressor stages. The low pressure air may be used to aid in cooling components of the engine as well.

The gas turbine 10 is axis-symmetrical about engine axis 26 or shaft 24 so that various engine components rotate thereabout. The axis-symmetrical shaft 24 extends through the turbine engine forward end into an aft end and is journaled by bearings along the length of the shaft structure. The shaft rotates about a centerline 26 of the engine 10. The shaft 24 may be hollow to allow rotation of a low pressure turbine shaft 28 therein and independent of the shaft 24 rotation. Shafts 28 also may rotate about the centerline axis 26 of the engine. During operation the shaft 28 rotates along with other structures connected to the shaft such as the rotor assemblies of the turbine in order to create power or thrust for various types of turbines used in power and industrial or aviation areas of use.

At the forward end 12 of the engine 10, forward of the turbo fan blades 18 is a nose cone, also referred to as a spinner 30. The spinner 30 is generally attached to a fan hub in a variety of fashions including but not limited to a number of circumferentially spaced bolts. The spinner 30 is utilized to provide a smooth flow of air to the core or radially inner portions of the fan 18. Smoothing of the airflow increases efficiency of the engine 10 and therefore improves performance not only of the fan 18, but of downstream components as well. For example, the spinner 30 shape may reduce drag, correct velocity profile into the core, reduce turbulence into the core, as well as provide a means for shedding ice and/or deflect foreign objects toward the fan/bypass ducts rather than allowing passage through the core, which can damage engine components. The parts or components such as the spinner 30 and aft support ring 50 of the instant disclosure are formed of hybrid polymer matrix composites, wherein a first portion is formed of a first fiber type and a second portion is formed of a second fiber type. One of the first fiber type and second fiber type is used to form less complex shapes that have higher loading while the other of the first and second fiber types is used to form more complex shapes that have lighter loading. Despite the two fiber types, the aircraft components being formed are monolithic.

Composite materials generally comprise a fibrous reinforcement material embedded in matrix material, such as polymer or ceramic material. The reinforcement material serves as a load-bearing constituent of the composite material, while the matrix of a composite material serves to bind the fibers together, and also acts as the medium by which an externally applied stress is transmitted and distributed to the fibers. Many polymer matrix composite (PMC) materials are fabricated with the use of prepreg, which is a fabric or unidirectional tape that is impregnated with resin. Multiple layers of prepreg are stacked to the proper thickness and orientation for the part, and then the resin is cured and solidified to render a fiber reinforeced composite part. Resins for matrix materials of PMCs can be generally classified as thermosets or thermoplastics. Thermoplastic resins are generally categorized as polymers that can be repeatedly softened and flowed when heated and hardened when sufficiently cooled due to physical rather than chemical changes. Notable example classes of thermosplastic resins include nylons, thermoplastic polyesters, polyaryletherketones, and polycarbonate resins. Specific example of high performance thermoplastic resins that have been contemplated for use in aerospace applications include, polyetheretherketone (PEEK), polyetherketoneketone (PEKK), polyetherimide (PEI) and polyphenylene sulfide (PPS). In contrast, once fully cured into a hard rigid solid, thermoset resins do not undergo significant softening when heated, but instead thermally decompose when sufficiently heated. Notable examples of thermoset resins include epoxy, bismaleimide (BMI), and polyimide resins.

A variety of fibrous reinforcement materials have been used in PMCs, for example, carbon (e.g., AS4), glass (e.g., S2), polymer (e.g., Kevlar®), ceramic (e.g. Nextel®) and metal fibers. Fibrous reinforcement materials can be used in the form of relatively short chopped fibers, generally less than two inches in length, and more preferably less than one inch, or long continuous fibers, the latter of which are often used to produce a woven fabric or unidirectional tape. PMC materials can be produced by dispersing dry fibers into a mold, and then flowing matrix material around the reinforcement fibers, or by using prepreg as previously described.

Whether a PMC material is suitable for a given application depends on its matrix and reinforcement materials, the requirements of the particular application, and the feasibility of fabricating a PMC article having the required geometry. Due to their considerable potential for weight savings, various applications have been explored for PMCs in aircraft gas turbine engines. However, a challenge has been the identification of material systems that have acceptable properties yet can be produced by manufacturing methods to yield a cost-effective PMC component. In particular, it is well known that aircraft engine applications have high performance mechanical requirements, for example, strength and fatigue properties (necessitated by vibrations in the engine environment), as well as high temperature properties, chemical/fluid resistance, etc. Though considerable weight savings could be realized by fabricating engine parts from PMC materials, performance requirements as well as the size and complexity of such components have complicated the ability to produce components from these materials.

Another complication is the type of reinforcement system required by PMC materials in aircraft engine applications. Generally, to achieve the mechanical properties required for aircraft engine applications, parts would require the use of continuous fiber-reinforced PMC materials to achieve the high performance mechanical requirements (particularly strength and fatigue properties) dictated by aircraft engine applications. However, the manufacturing processes involved in the fabrication of continuous fiber reinforcement composite parts further complicate the ability to produce structures that have complex shapes. On the other hand, chopped fiber reinforcement systems, whether in thermoplastic or thermoset resin matrix, are not ideal solutions for highly loaded parts due to their lower mechanical performance. However, it is possible to fabricate complex-shaped parts with chopped fiber material solutions with net-shaped molding methods, and therefore these material systems can be used for lightly-loaded secondary structures and non-structural engine components.

As engine performance continues to be pushed to limits, it is desirable to have parts of complex geometries that are capable of being highly loaded to aid or improve such performance. Many times, these complex geometries are non-structural features that help with, for example, aerodynamic performance. Therefore, taking a hybrid approach a monolithic part is provided with hybrid fiber reinforcement to achieve structural loading yet providing for the complex shaped (lightly loaded) features, for example aero-features.

Referring now to FIG. 2, a part is shown which is used toward the fan end of the engine 10. Although components of the fan end of the engine are shown and described, other engine parts may be formed using the design shown and described herein and the exemplary parts should not be considered limiting. The spinner 30 is generally formed of a conical shape formed of a sidewall 32 which is generally continuous. The conical shaped sidewall 32 tapers from a first end 36 to a larger second end 38. The sidewall 32 may be linear moving from the larger end of the cone to the smaller end of the cone. Alternatively, the sidewall 32 may be curvilinear as shown. The spinner 30 is symmetrical about the axis 34, shown in broken line. The spinner 30 is generally hollow to reduce weight and is capable of receiving bolts, fixtures or other components of the fan hub (not shown).

At the forward end 12 of the engine 10 (FIG. 1), the engine temperatures are lower which permits the use of PMC materials for the spinner 30. The spinner 30 has significant loading requirements. Design characteristics include, for example, aerodynamic loading, high speed revolution fatigue and foreign object strikes. Accordingly, the spinner 30 is formed of polymer matrix composite and more specifically may be formed of continuous fiber polymer composite material. The continuous fiber polymer composites may provide the conical or parabolic conical shape desired for the spinner 30 as these shapes are easily formed with polymer composite materials. According to this design, the spinner weight may exhibit significant weight reduction, for example between 5 to 20 pounds over a metallic design, depending on the engine type. For example, the shape of the cone spinner 30 is generally consistent without sharp changes from the forward end 36 of the cone toward the aft end 38. Accordingly, much of the spinner 30 may be formed by laying up continuous fiber portions that are in a fabric, unidirectional tape, or braided architecture. Each of the continuous fiber portions may be rotated to a preselected angle layer by layer to achieve the strength required for the part.

At the aft end 38 of the spinner 30 are a plurality of circumferentially spaced aero-features or second features 40. These second features 40 extend from the surface 32 of the spinner 30 and provide a geometry which is complicated to form by continuous fiber composite fabrication methods. For example, the aero-features 40 extend from the surface 32 at various angles and may be of varying thickness making difficult the use of continuous fiber composites as well as known techniques for manufacturing with such continuous fiber composite. Accordingly, in order to form a monolithic part, such as the spinner 30 depicted, the aero-features 40 are formed of thermoplastic polymer unidirectional tape that has been chopped to a short fiber length. The thermoplastic polymer used in the chopped fiber unidirectional tape, is the same thermoplastic polymer that is used in the continuous fiber portion of the part, for example, spinner 30. This allows the first fibers, second fibers and polymer to be fabricated as the monolithic part depicted. The cone sidewall or body 32 may be formed of, for non-limiting example, unidirectional prepreg, woven fabric prepreg, a braided prepreg, or a dry reinforcement fiber with filaments or fibers of thermoplastic polymer. For example the continuous fiber material may be continuous fibers of individual fibers or fiber tows arranged parallel (unidirectional) with the matrix material, or individual fibers or fiber tows arranged to have multiple different orientations (e.g., multiple layers of unidirectional fibers or fiber tows to form bi-axial or tri-axial architecture) within the matrix material, or individual fibers or fiber tows, woven to form a mesh or fabric within the matrix material. The fibers, tows, braids, meshes or fabrics can be arranged to define a single ply within the PMC or any suitable number of plies. Particularly suitable thermoplastic matrix materials include PEEK, PEKK, PEI and PPS and particularly suitable continuous fiber reinforcement materials include carbon, glass polymer, ceramic and metal fibers. Suitable fiber content may be at least 35 percent by volume and not more than 75 percent by volume, with a preferred range believed to be about 50 to about 65 percent by volume.

According to one embodiment, the PMC material is defined in part by prepreg, which is a reinforcement material preimpregnated with a matrix material, such as thermoplastic resin desired for the matrix material. Non-limiting examples of processed for producing thermoplastic prepregs include hot melt prepregging in which the fiber reinforcement material is drawn through the molten bath of resin, and powder prepregging in which a resin is deposited onto the fiber reinforcement material (for example electrostatically) and then adhered to the fiber (for example, in an over or with the assistance of heated rollers). The prepregs can be in the form of unidirectional tapes or woven fabrics, which are then stacked on top of one another to create the number of stacked plies desired for the part. According to an alternative option, instead of using a prepreg, with the use of thermoplastic polymers it is possible to have a woven fabric that has, for example dry carbon fiber woven together with thermoplastic polymer fibers or filaments. Non-prepreg braided architectures can be made in a similar fashion. With this approach, it is possible to tailor the fiber volume of the part by dictating the relative concentrations of the thermoplastic fibers and reinforcement fibers that have been woven or braided together. Additionally, different types of reinforcement fibers can be braided or woven together in various concentrations to tailor the properties of the part. For example, glass fiber, carbon fiber, and thermoplastic fiber could all be woven together in various concentrations to tailor the properties of the part. The carbon fiber provides the strength of the system, the glass may be incorporated to enhance the impact properties, which is a design characteristic for parts located near the inlet of the engine, and the thermoplastic fibers are the matrix that will be flowed to bind the reinforcement fibers.

The ply stack may next undergo a consolidation operation, in which heat and pressure are applied to the ply stack to flow the resin and consolidate the ply stack into the part. In addition to creating parts using prepreg, an alternative approach is to lay-up dry fabric in a suitably shaped mold cavity and then infuse the dray fabric with molten resin.

According to the instant embodiment, due to its shape, the spinner cone 30 continuous fiber preform architecture is loaded into a compression mold. Within this mold are cavities corresponding to the shape of features 40 wherein chopped fiber unidirectional tape prepreg flakes are loaded into the mold cavities to form the aero-features 40. The combination of continuous and chopped fibers are then molded into a final part rendering, for example, a monolithic spinner cone 30, including the continuous fiber and chopped fiber sections.

Additionally, the part 30 may be machined if necessary such as by conventional machining, waterjet cutting, and laser cutting techniques. For example, the part 30 may be formed to include slots, holes or other features with which a component or assembly structure, etc., could be mounted to gas turbine engine through the use of conventional mechanical fasteners and/or attachment mechanisms. Additionally, metallic features may be co-molded with the part 30 to enable more robust mechanical fastening. Non-limiting examples include co-molded metallic bushings, co-molded metallic attachment rings, and co-molded metallic threaded inserts. Another advantage of thermoplastic composite materials is that they can undergo various joining processes including, but not limited to, infrared (IR) welding, resistive implant welding, ultrasonic welding, and vibration welding.

As a result of the construction, a load-bearing part is formed which benefits from weight savings but also has requisite capability and characteristics for withstanding mechanical and environmental conditions associated with aircraft engines. Additionally, a monolithic hybrid composite structure may be fabricated which can withstand high loadings yet contain complex secondary features.

Referring now to FIG. 3, an isometric view of a second part or component 50 is depicted which is according to the described embodiment capable of being formed of a first portion formed of continuous fiber and a second portion of more complex geometry formed of chopped fiber. According to the instant embodiment, the part 50 is an aft support ring. The aft support ring 50 is generally circular in cross-section and includes a body or first surface 52 extending aft from a forward flange 54. The flange 54 includes a plurality of fastening apertures 56 through which the aft support ring 50 may be connected to the aft end 38 of the spinner 30. The arm 52 includes a plurality of flow path scallops 58 which aid to improve aerodynamic flow of the air leaving the spinner 30 and moving across the aft support ring 50 from the forward end toward the aft end of the ring 50.

Referring now to FIG. 4, a section view of the aft support ring 50 is depicted in an assembly. At the forward end of the aft support ring 50 is a portion of the spinner 30 which is fastened to the flange 54 through aperture 56. The aft support ring 50 includes the flange 54, curved arm 60, the body or arm 62, a lug 64 and the flow path scallop 58. The lug 64 and the scallop 58 define the interface wherein the continuous fibers reinforcement section transitions to the chopped fiber reinforcement section.

In manufacturing, the flange 54, curved arm 60, the body 62 and lug 64 may be formed of continuous fiber reinforcement to carry the high loads that the part is subjected to. The preformed architecture of the flange 54, curved arm 60, body 62 and lug 64 may be formed of dry carbon fiber and thermoplastic polymer fibers braided into the preformed structure. Further, for example, glass fibers may be added to the preform to improve impact characteristics. The flow path scallops 58 are formed on the lug 64 and the body 62 with chopped fiber unitape prepreg flakes. Such continuous fiber preform is loaded into a compression mold and the chopped fiber unitape prepreg flakes are loaded into flow path scallop cavities in the mold to define the flow path scallop 58 shape. The compression molding is a non-limiting example as other methods may be utilized. For example an autoclave may be an alternative method. These chopped fiber unitape flakes may be less than 1″×1″, for example, ½″×½″, in size although alternate shapes and sizes may be utilized. The combinations of continuous fiber preform architecture and chopped fiber unitape are compression molded into the final shape of the part, rendering a monolithic part with both continuous fiber and chopped fiber sections. During the compression molding process, the thermoplastic polymer in the continuous fiber reinforced section and chopped fiber reinforced section flow together to create a monolithic part that is made from one polymer type, while having the two fiber types.

As shown in FIG. 5, a schematic section view is depicted for the aft support ring 50. The scallop 58 is curved from the surface 52 to an upper height 59. The scallop 58 is disposed above the lug 64. As shown at the interface of the lug 64 and the scallop 58, the different cross-hatchings depict the interface or joining area of the continuous fibers 66 of the lug 64 and the chopped fibers 68. Between the fibers 66, 68 are where the thermoplastic resin 70 flows to bind the continuous fibers 66 to the chopped fibers 68. Also the thermoplastic 70 flows between like fiber types to provide for the monolithic part.

The present hybrid continuous fiber-chopped fiber polymer composite provides various benefits. The continuous fibers and chopped fibers utilize the same resin or polymer, thereby eliminating compatibility issues when transitioning from the continuous fiber reinforced section to the chopped fiber reinforced section. The component precludes the need to make aero features separately and a subsequent joining process step to bond the two portions of the component, for example by adhesive bonding or mechanical fastening. Since the parts or structures may be formed as a single monolithic structure, in certain situations this will enable a reduction in part count and eliminate surface preparation that would be needed in joining the two components in alternative designs. Further, the hybrid monolithic polymer composite may be formed to improve impact resistance at complex secondary feature locations, as the strength of the thermoplastic polymer is typically higher than that of an adhesive bond between two different materials that would be used in alternative designs. Additionally, the polymer composite may be tailored to the desired impact properties by incorporating mixtures of carbon and/or glass fiber into the continuous fiber preform. Even further, an erosion protection layer may be deposited on an outermost surface of any of the exemplary parts described.

Related to the reduction of parts described above, the hybrid polymer composite enables co-molding of metallic features as well as thermoplastic welding and these various benefits and applications may be utilized with numerous parts including but not limited to the spinner 30 and the aft support ring 50 described herein. For example, metallic inserts may be utilized to aid strength and provide an advantage to eliminate mechanical fastening directly on the composite parts, for example, at the interface of the spinner 30 and aft support ring 50 (FIG. 4).

The foregoing description of structures and methods has been presented for purposes of illustration. It is not intended to be exhaustive or to limit the structures and methods to the precise forms and/or steps disclosed, and obviously many modifications and variations are possible in light of the above teaching. Features described herein may be combined in any combination. Steps of a method described herein may be performed in any sequence that is physically possible. It is understood that while certain forms of composite structures have been illustrated and described, it is not limited thereto and instead will only be limited by the claims, appended hereto.

While multiple inventive embodiments have been described and illustrated herein, those of ordinary skill in the art will readily envision a variety of other means and/or structures for performing the function and/or obtaining the results and/or one or more of the advantages described herein, and each of such variations and/or modifications is deemed to be within the scope of the embodiments described herein. More generally, those skilled in the art will readily appreciate that all parameters, dimensions, materials, and configurations described herein are meant to be exemplary and that the actual parameters, dimensions, materials, and/or configurations will depend upon the specific application or applications for which the inventive teachings is/are used. Those skilled in the art will recognize, or be able to ascertain using no more than routine experimentation, many equivalents to the specific inventive embodiments described herein. It is, therefore, to be understood that the foregoing embodiments are presented by way of example only and that, within the scope of the appended claims and equivalents thereto, inventive embodiments may be practiced otherwise than as specifically described and claimed. Inventive embodiments of the present disclosure are directed to each individual feature, system, article, material, kit, and/or method described herein. In addition, any combination of two or more such features, systems, articles, materials, kits, and/or methods, if such features, systems, articles, materials, kits, and/or methods are not mutually inconsistent, is included within the inventive scope of the present disclosure.

Examples are used to disclose the embodiments, including the best mode, and also to enable any person skilled in the art to practice the apparatus and/or method, including making and using any devices or systems and performing any incorporated methods. These examples are not intended to be exhaustive or to limit the disclosure to the precise steps and/or forms disclosed, and many modifications and variations are possible in light of the above teaching. Features described herein may be combined in any combination. Steps of a method described herein may be performed in any sequence that is physically possible.

All definitions, as defined and used herein, should be understood to control over dictionary definitions, definitions in documents incorporated by reference, and/or ordinary meanings of the defined terms. The indefinite articles “a” and “an,” as used herein in the specification and in the claims, unless clearly indicated to the contrary, should be understood to mean “at least one.” The phrase “and/or,” as used herein in the specification and in the claims, should be understood to mean “either or both” of the elements so conjoined, i.e., elements that are conjunctively present in some cases and disjunctively present in other cases.

It should also be understood that, unless clearly indicated to the contrary, in any methods claimed herein that include more than one step or act, the order of the steps or acts of the method is not necessarily limited to the order in which the steps or acts of the method are recited.

In the claims, as well as in the specification above, all transitional phrases such as “comprising,” “including,” “carrying,” “having,” “containing,” “involving,” “holding,” “composed of,” and the like are to be understood to be open-ended, i.e., to mean including but not limited to. Only the transitional phrases “consisting of” and “consisting essentially of” shall be closed or semi-closed transitional phrases, respectively, as set forth in the United States Patent Office Manual of Patent Examining Procedures, Section 2111.03. 

What is claimed is:
 1. An aircraft engine component, comprising: a monolithic composite body; said body having: a continuous fiber portion; a chopped fiber portion; a thermoplastic polymer contained in both said continuous fiber portion and said chopped fiber portion and between said continuous and chopped fiber portions.
 2. The aircraft engine component of claim 1, wherein said thermoplastic polymer is one of PEEK, PPS, PEKK, and PEI.
 3. The aircraft engine component of claim 1, wherein the fiber is one of carbon fiber, glass fiber and a mixture of said carbon fiber and said glass fiber.
 4. The aircraft engine component of claim 1, wherein said continuous fiber is one a braid, woven fiber and a unidirectional tape.
 5. The aircraft engine component of claim 1, wherein said chopped fiber constituent is formed of a unidirectional preimpregnated tape.
 6. The aircraft engine component of claim 5, wherein said chopped fiber length is less than one inch (1″).
 7. The aircraft engine component of claim 1 further comprising one or more co-molded metallic features.
 8. The aircraft engine component of claim 7, said co-molded metallic features being one of a flange, bushing and threaded insert.
 9. The aircraft engine component of claim 1, said continuous fiber of said composite body carrying a hoop load.
 10. The aircraft engine component of claim 1, wherein said continuous fiber is a braided preform that contains at least one of fibers of carbon, glass and thermoplastic.
 11. The aircraft engine component of claim 10 wherein said at least one of fibers is dry carbon fiber and thermoplastic fiber.
 12. The aircraft engine component of claim 10 wherein said at least one of fibers is dry carbon fiber, glass fiber and thermoplastic fiber.
 13. The aircraft engine component of claim 1, wherein said chopped fiber portion fails to carry a structural load.
 14. The aircraft engine component of claim 1, wherein said chopped fiber portion is adjacent to an engine air flowpath.
 15. The aircraft engine component of claim 14 further comprising an erosion protection layer on an outermost surface of said chopped fiber portion.
 16. The aircraft engine component of claim 1, said component being a rotating part.
 17. The aircraft engine component of claim 16, said rotating part being a spinner nose cone.
 18. The aircraft engine component of claim 16, said rotating part being a spinner support ring. 